1. Field
The present invention relates to gas turbine engines, and more specifically to a turbine blade with multiple internal cooling air circuits.
2. Description of the Related Art
In an industrial gas turbine engine, hot compressed gas is produced. The hot gas flow is passed through a turbine and expands to produce mechanical work used to drive an electric generator for power production. The turbine generally includes multiple stages of stator vanes and rotor blades to convert the energy from the hot gas flow into mechanical energy that drives the rotor shaft of the engine. Turbine inlet temperature is limited to the material properties and cooling capabilities of the turbine parts. This is especially important for first stages of turbine vanes and blades since these airfoils are exposed to the hottest gas flow in the system.
A combustion system receives air from a compressor and raises it to a high energy level by mixing in fuel and burning the mixture, after which products of the combustor are expanded through the turbine.
Since the turbine blades are exposed to the hot gas flow discharged from combustors within the combustion system, cooling methods are used to obtain a useful design life cycle for the turbine blade. Blade cooling is accomplished by extracting a portion of the cooler compressed air from the compressor and directing it to the turbine section, thereby bypassing the combustors. After introduction into the turbine section, this cooling air flows through passages or channels formed in the airfoil portions of the blades. The blade tip and the trailing edge of the blade are the most challenging locations in cooling.
The turbine second row blade is typically larger than the first row blade and has more surface area to cool. The second row blade is exposed to a lower gas temperature than the first row blade and therefore needs to allow for the use of less amounts of cooling air for better turbine efficiency.
In order to allow for higher temperatures, turbine blade designers have proposed several complex internal blade cooling circuits to maximize the blade cooling through the use of convection cooling, impingement cooling and film cooling of the blades. Conventionally, the focus of cooling improvement has been with the first row blade for more impact to turbine efficiency.
FIGS. 1 and 2 show a prior art turbine blade with two cooling circuit designs. The cooling circuit designs include a forward and an aft circuits. The forward blade cooling circuit includes a channel that exits to an axial tip cooling channel. The aft blade cooling circuit includes a first pass cooling channel, a second pass cooling channel, and a third pass cooling channel. The cooling circuit flow from the mid-chord aft ward towards a trailing edge of the blade.
FIGS. 1 and 2 show a turbine blade with two cooling circuits. The leading edge circuit enters from the leading edge and flowing aft with a 90 degree turn at the tip and exits the blade through a tip axial cooling channel. The trailing edge circuit enters from the mid-chord and flowing aft with two 180 degree turns at the tip and the root and exits through trailing edge pin banks and trailing edge exit holes. The third leg or pass cooling channel of the trailing edge circuit may include a row of exit cooling holes or slots to discharge the cooling air from the serpentine flow circuits axially out from the blade.
While this design provides good cooling to the majority of the airfoil, the blade tip section is much hotter than the other portions of the airfoil.